A turbomachine combustion chamber is generally annular in shape, centered on an axis X that corresponds to the axis of rotation of the rotor of the turbomachine. The combustion chamber has two annular walls (or shrouds) disposed coaxially about the axis X, and a chamber end wall arranged between said shrouds in the upstream region of said chamber. Said chamber walls define the combustion enclosure of the chamber.
A plurality of fuel injection systems are fastened to the end wall of the chamber (in orifices provided for this purpose through the chamber end wall), and they are distributed regularly around the axis X. A fuel injector is connected to each injection system.
An injection system generally comprises a central hole for receiving the free end of a fuel injector, and a plurality of air admission passages disposed concentrically around said central hole. These air admission passages generally serve to swirl air, i.e. they are constituted by annular passages having series of vanes extending therein for imparting rotary motion to the air passing therethrough. The air passing through these admission passages comes from the diffuser of the turbomachine, which diffuser is situated upstream from the combustion chamber. This air is mixed with fuel delivered by said injector so as to form an air/fuel mixture that is burnt in the combustion chamber.
The injector is a part that passes through the “combustion chamber module” of the turbomachine from the outer casing of said module to the end wall of the chamber. The injector comprises the fuel feed pipe(s) (forming part of a fuel manifold) that serve(s) to convey fuel for injection into the chamber. The fuel is expelled from the free end, or “nose”, of the injector.
The invention relates to a fuel injector and not to the injection system to which the injector is connected.
The conventional process for designing and optimizing an airplane turbojet combustion chamber seeks to reconcile the expected operating performance (typically: fuel efficiency, stability range, ignition and re-ignition range, combustion area lifetime, temperature distribution at the outlet from the combustion area) depending on the intended mission of the airplane, with reducing polluting emissions (NOx, CO, UHC, soot).
A known solution for reducing polluting emissions, in particular those involving nitrogen oxides (of the NOx type) is to ensure that the combustion flame is in the presence of an air/fuel mixture that is rich or lean. For example, a leaner air/fuel mixture can be obtained for the combustion flame by increasing the flow rate of the air delivered to the combustion (mainly the flow rate through the end wall of the chamber).
Experience shows that this solution based on a lean mixture is effective in reducing NOx. Nevertheless, when attempts are made to take this solution to its potential maximum in terms of reducing polluting emissions (for a combustion area of fixed shape and thus of constant air distribution), it suffers from the following drawback: the stability of the combustion flame is affected (i.e. the vulnerability of the flame to being blown out is increased), in particular at low operating speeds of the turbojet, such that idling stages can no longer be obtained for the engine; the combustion efficiency at intermediate operating speeds is reduced; when the airplane is on the ground, it is more difficult to ignite the flame; and re-igniting the flame when the airplane is at altitude is also more difficult.